Multi lift aircraft control system

ABSTRACT

A single load is suspended from vertical cables extending beneath each of two direct lift type of aircraft, the cables being separated by a spreader bar. A trailing, slave aircraft flys behind, above, and to one side of a leading, master aircraft, with the spreader bar extending at an angle to the heading of the aircraft, in a plane below them. The lead pilot control provides overall, or outer loop control of the lead aircraft, and the controls of the lead aircraft are electromechanically coupled to operate the controls of the trailing aircraft, whereby the lead pilot controls the outer loop of both aircraft. The inner or stability loop of the controls of each aircraft are modified so as to utilize inputs from the inertial system of the related aircraft and inputs relating to the manner in which the load extends below the aircraft, to both stabilize the load with respect to each aircraft, and to maintain the trail aircraft in a proper position with respect to the lead aircraft and the load. Specifically, the cables from each aircraft supporting the spreader bar are maintained substantially perpendicular with respect to the spreader bar thereby to maintain substantially equal loading of each aircraft; pitch and roll inertial sensors in each aircraft, and the angle of each cable with respect to its aircraft in both the pitch and roll directions, are utilized in the respective aircraft to stabilize that aircraft with respect to the load, thereby to prevent wiggling of the load; the heading of the load itself and of each aircraft, together with desired formation inputs, are utilized to control the heading of the trailing aircraft, and to provide pitch and roll inputs to control relative position of the trailing aircraft with respect to the leading aircraft so as to maintain a desired formation in flight.

United States Patent 1191 Maciolek et a1.

[ MULTI-LIFT AIRCRAFT CONTROL SYSTEM [75] Inventors: Joseph R. Maciolek,Newtown; John J. Wallace, Winsted, both of Conn.

[73] Assignee: United Aircraft Corporation, East Hartford, Conn,

[22] Filed: Aug. 2, 1971 [21] App]. No.: 168,219

52 US. Cl 244 2, 244/17.13, 244/77 D 51 1m. 0.... B64c 37/02, B64c11/44, B640 27/70 [58] Field of Search 235/1502, 183, 197;

244/2, 77 R, 77 D, 77 V, 84, 17.13; 307/229; 318/489, 584, 585

Primary Examiner-Milton Buchler Assistant Examiner-Stephen G. KuninAttorneyMelvin Pearson Williams [57] ABSTRACT A single load is suspendedfrom vertical cables extending beneath each of two direct lift type ofaircraft, the

July 17, 1973 cables being separated by a spreader bar. A trailing,slave aircraft flys behind, above, and to one side of a leading, masteraircraft, with the spreader bar extending at an angle to the heading ofthe aircraft, in a plane below them. The lead pilot control providesoverall, or outer loop control of the lead aircraft, and the controls ofthe lead aircraft are electromechanically coupled to operate thecontrols of the trailing aircraft, whereby the lead pilot controls theouter loop of both aircraft, The inner or stability loop of the controlsof each aircraft are modified so as to utilize inputs from the inertialsystem of the related aircraft and inputs relating to the manner inwhich the load extends below the aircraft, to both stabilize the loadwith respect to each aircraft, and to maintain the trail aircraft in aproper position with respect to the lead aircraft and the load.Specifically, the cables from each aircraft supporting the spreader barare maintained substantially perpendicular with respect to the spreaderbar thereby to maintain substantially equal loading of each aircraft;pitch and roll inertial sensors in each aircraft, and the angle of eachcable with respect to its aircraft in both the pitch and rolldirections, are utilized in the respective aircraft to stabilize thataircraft with respect to the load, thereby to prevent wiggling of theload; the heading of the load itself and of each aircraft, together withdesired formation inputs, are utilized to control the heading of thetrailing aircraft, and to provide pitch and roll 12 Claims, 18 DrawingFigures re /2 Min/0,42 27 PATENIELJUL 1 H913 SHEET 5 OF 8 PATENTEU JUL 17 ms SHEET 8 0F 8 MULTI-LIFT AIRCRAFT CONTROL SYSTEM BACKGROUND OF THEINVENTION 1. Field of Invention This invention relates to controlsystems for aircraft, and more particularly to a multi-lift,master/slave aircraft control system.

2. Description of the Prior Art Utilization of direct lift-type ofaircraft, such as helicopters, has recently been expanded to includeheavy lift operations, in the nature of a crane in the sky. Naturally,the load that can safely be supported by an individual aircraftapproaches limitations asymptotically in the sense that lifting of agreater load requires stronger support mechanisms, larger engines, morefuel, and a larger aircraft in general; the weight of the aircrafttherefore increases in proportion to the weight of the load which it isto lift.

One solution to this problem which has been proposed in the prior art isto utilize a plurality of aircraft, in common, to jointly support aload. One such system known to the art simply provides for thesuspension of a spreader bar between two cables extending below twocompletely independent aircraft, the load being suspended from cablesattached to the ends of the spreader bar. In this system, each aircraftis completely independently controlled by the pilots thereof, with nocommonality of control between the aircraft, other than visualobservations of the relative positions and attitudes of each aircraftwith respect to each other and with respect to the load. However, it hasbeen found that such systems readily attain, even in a simple hoveringmaneuver, shift in the amount of load supported by each aircraft ofpercent of the load or more, on a regular basis. When the aircraft areturning, accelerating, decelerating, or most particularly gaining orlosing altitude, the amount of load shifted from one aircraft to theother can easily approach 50 percent. When it is considered that eitheraircraft could carry a load half as great all by itself, and the needfor a dual lift system is to carry loads in excess of those which can besustained by a single aircraft, it is quite obvious that a shift of anysubstantial fraction of the load (in excess of 50 percent of the load)to a single aircraft is unsatisfactory.

Another type of system known to the art is the interconnection of two ormore aircraft through a rigid load carrying member, pilot control overone aircraft causing, by mechanical interconnection, common control ofthe other aircraft. However, this system provides severe limitations onthe independent movement of each of the aircraft, rendering themaneuvering thereof extremely difficult and dangerous, and providingsevere limitations upon the rates of change of position which may beachieved through maneuvering of the aircraft. Additionally, since theoperation of any one aircraft causes an identical input in the otheraircraft, it is impossible to maintain relative formation while changingheading or altitude; instead, all aircraft must exhibit the sameelevation and heading changes at the same moment. This means that onlyflank type motions can be achieved by the formation, and no column-typemotions can be achieved by the formation. The relative position of eachaircraft in the formation is changed as a result of any changes inheading of the group. In addition, since no two aircraft are exactlyalike, the response to a given control input will not be exactly alike;

therefore each aircraft is liable to exhibit a loading effect as aresult of a different response to a given command in any of the otheraircraft. This results in further dangerous and possibly catastrophicconditions, rendering the utilization of such a system inadvisable.

SUMMARY OF INVENTION The object of the present invention is to providean improved multiple lift aircraft control system.

In accordance with one aspect of the present invention, a plurality ofaircraft suspend a load by mechanically-spread cables extending beneatheach aircraft; pilot control over one of the aircraft not only providesouter loop (or overall) control over the aircraft, but throughelectro-mechanical coupling provides control over the outer loop of theother aircraft. In further ac cord with the present invention, the innerloop (or stability) control of each aircraft is modified to providestability of each aircraft with respect to the load and each other.

In accordance with another aspect of the present invention, in a systemin which a load is supported by mechanically spread cables extendingbeneath each of a plurality of direct lift type of aircraft, thedifference of a pair of angles between the mechanical cable spreadingmeans and the cable relating to a respective pair of the aircraft isutilized to control direct lift of one of the aircraft, thereby tendingto maintain the cable load of that aircraft substantially normal to thetotal force field exerted on the spreader means as a result of gravityand accelerations due to speed up, slow down or turns.

In accordance with still another aspect of the present invention,relative heading of the master aircraft and the load is utilized tocontrol position of a slave aircraft in a multi aircraft load liftingsystem, thereby tending to maintain a desired formation of thesupporting aircraft with respect to the load.

The present invention provides a relatively simple, stable and safesystem for controlling a plurality of aircraft with respect to a commonload. The invention provides not only for control of slave aircraft inresponse to maneuvering of a master aircraft, but also provides loadstability control over all of the aircraft. The invention permitsindependent (though related) maneuvering of each of the aircraft whilemaintaining a stable load, and permits maneuvering all of the aircraftin a relative formation in a wide range of maneuvers.

Aspects of the invention may be used separately, in low cost,partially-automated control systems.

Other objects, features and advantages of the present invention willbecome more apparent in the light of the following detailed descriptionof preferred embodiments thereof, as illustrated in the accompanyingdrawmg.

BRIEF DESCRIPTION OF THE DRAWING FIG. l is a simplified side viewillustration of a dual helicopter lift operation while hovering;

FIG. 2 is a simplified front view illustration of a dual helicopter liftoperation while hovering;

FIG. 3 is a simplified top view illustration of a dual helicopter liftoperation while hovering;

FIG. 4 is a simplified side view illustration of a dual helicopter liftoperation in correctly-executed forward flight;

FIG. 5 is a simplified side view illustration of a dual helicopter liftoperation in incorrectly executed forward flight;

FIG. 6 is a simplified front view illustration of a dual lift helicopteroperation correctly executing a turn to starboard;

FIG. 7 is a simplified front view illustration of a dual lift helicopteroperation incorrectly executing a turn to starboard;

FIG. 8 is a schematic block diagram of a direct lift control which maybe used in a dual lift aircraft control system;

FIG. 9 is a partial schematic block diagram of an alternative anglesensor which may be utilized in the embodiment of FIG. 8 in accordancewith one embodiment of the invention;

FIG. 10 isa simplified front view illustration of a dual lift helicopteroperation with the aircraft located too close together laterally;

FIGQl I is a simplified front view illustration of a dual lifthelicopter operation with incorrect relative elevation of the aircraft;

FIG. 12 is a simplified, partial schematic-block diagram of anotherembodiment of an angle sensor which may be utilized in the embodiment ofFIG. 8, in accordance with the invention;

FIG. 13 is a schematic block diagram of lateral (or roll) cyclic pitchcontrols in accordance with one embodiment of the present invention;

FIG. 14 is a schematic block diagram of longitudinal (or pitch) cyclicpitch controls in accordance with one embodiment of the presentinvention;

FIG. 15 is a schematic block diagram of heading controls which may beused in an embodiment of the present invention;

FIG. 16 is a schematic block diagram of position controls used inconjunction with the lateral and longitudinal cyclic controlsillustrated in FIGS. 13 and 14, in accordance with the invention;

FIG. 17 is a simplified schematic block diagram illustrating concepts ofsingle aircraft control known to the prior art; and I FIG. 18 is asimplified schematic block diagram illustrating dual liftaircraftcontrol concepts in accordance with an embodiment of the presentinvention.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to FIGS. 1-3, apair of helicopters 20, 22 each support a respective cable 24, 26connected to opposite ends 28, 30 of a spreader bar 32, which supports,by means of equal length cables 34, 36, a load 38 to be shared by thehelicopters 20, 22. As is viewed in FIGS. 1-3, the helicopters 20, 22are shown hovering with no velocity in any direction). As seen therein,the lead helicopter is at an altitude slightly less than the trailinghelicopter 22. This is to facilitate a comfortable viewing angle for thepilot of the trailing helicopter 22. Additionally, a standard formationincludes the trailing aircraft at a higher altitude than the leadingaircraft. Also, it is essential that the trailing aircraft not be at toolow an altitude relative to the leading aircraft due to the effects ofdownwash from the leading aircraft which could alter the stable flightpattern of the trailing aircraft.

As is illustrated in FIGS. 1-3, the cables 24, 26 are at right angleswith respect to the bar 32. This is essential in order to equally sharethe load between the two helicopters. Regardless of the maneuvering ofthe helicopters, this characteristic continues to be true: if the cables24, 26 are not at substantially right angles to the bar 32, then theload carried by each helicopter will be altered proportionately.

Referring to FIG. 4, the case of forward flight of the helicopters (incontrast with the hovering of FIGS. 1-3) requires that each helicopterbe tilted due to the flight characteristics of the helicopter, as iswell known, and due to drag forces opposing the motion of the entireload assembly 40 (including the cables 24, 26, 34, 36, the bar 32 andthe load 38) which drags behind the vertical projection of thehelicopters. During accelerated flight in a forward direction, the anglewould be great (perhaps as great as illustrated in FIG. 4); on the otherhand, during unaccelerated forward flight, the angle might be somewhatless than that illustrated in FIG. 4. Notice that the cables 24, 26 arestill at right angles with respect to the bar 32, so that the load isdivided equally between both helicopters (including the drag forceload). In FIG. 5 is illustrated what would happen if, instead ofmaintaining the cables at right angles to the bar 32, the bar 32 weremaintained in a horizontal position during forward flight. In FIG. 5, itis assumed that the helicopters are accelerating in forward direction.Due to the drag forces on the load assembly 40, it is seen that the netforce is in a direct line with the cable 24 supported by the helicopter20. That means that the helicopter 20 is supporting the entire loadexcept for the weight and drag of the cables 26 and 36, and half that ofthe bar 32. This illustrates a primary precept of the present invention:in order for two aircraft to maintain full maneuverability in flight,and equally share a load, it is necessary that supporting cablesextending to a spreader bar 32 be maintained substantially perpendicularto the bar at all times.

Conditions which obtain during a turn to the starboad are illustrated inFIGS. 6 and 7. In FIG. 6, a correct turn in accordance with theprinciples of the present invention is illustrated as maintaining thecables 24, 26 perpendicular to the bar 32. This evenly distributes theload force, including the centrifugal force necessary to accelerate inthe turn, between the two helicopters 20, 22. To the contrary, FIG. 7illustrates the condition which would obtain if the bar 32 weremaintained in a horizontal position as each helicopter tilted to makethe turn. This condition is substantially identical to that illustratedin FIG. 5; that is, the combined centrifugal force and gravity vector ofthe load 38 is very nearly in a straight line with the cable 24supported by the helicopter 20. Therefore, the helicopter 20 will becarrying nearly all of this load in direct proportion to the angle bywhich the load vector is off of the bisector of the bar 32. However, inan incorrectly executed turn as is shown in FIG. 7, the helicopter 22will none the less maintain the load of its own cables 22, 36 and halfof the bar 32. Of course, similar situations obtain with respect toturns to port and decelerations.

Consider momentarily operation of the helicopters 20, 22 under totalmanual control while attempting to maintain an even balance of loadbetween them. Positioning the two helicopters in a directionsubstantially perpendicular to the ends of the bar is relatively easydue to the fact that the load is pulling down on the cables and eachhelicopter is thereby tethered to the end of the bar 32. It is alsorelatively easy to maintain a relative azimuthal position (asillustrated in FIG. 3) within tolerable limits in the same fashion asthat well known type flying formation is maintained. However, tomaintain the bar level in its force field (including accelerations anddecelerations in forward velocity, and accelerations necessary in turns,in addition to gravitational force) it is necessary, as describedhereinbefore, to maintain the cables 24, 26 at right angles to the bar32 (not just vertically above the ends of the bar), which is difficultto calculate when the net force vector is not vertical. Also it isextremely difficult to control relative elevation or altitude undermanual operation, since it is extremely difficult to measure or visuallyestimate altitudes down to a few feet, as would be necessary to maintainproper orthogonality between the cables and the bar.

The desirable flight characteristics, which are the precepts of thepresent invention described with respect to FIGS. 1-7 hereinbefore, areachieved in accordance with one aspect of the present invention by meansof apparatus, the primary function of which is to maintain relativeflight positions of the helicopters 20, 22 with respect to each otherand with respect to the bar 32 such that the cables 24, 26 maintain aright angle relationship with the spreader 32.

Because of the flight parameters of a helicopter (that is, using cyclicpitch to tip the helicopter so that it is drawn in one direction oranother), the cables 24, 26 always extend in a direction substantiallynormal to the reference platform of the aircraft (the plane of thefuselage or substantially the plane of the rotor of the aircraft). Thisis as illustrated in FIGS. 4-7. If the trailing aircraft 22, as seen inFIG. 5, were to experience direct lift (which is achieved by collectivepitch in a helicopter), it would fly in the direction of its rotor asillustrated by the arrow 42 in FIG. 5; this would correct the geometryof this situation to that of FIG. 4, without any corrective action beingapplied to the lead aircraft 20. Similarly, if as illustrated in FIG. 7,the trailing aircraft 22 were provided with a direct lift input (asillustrated by the arrow 44), it would correct the situation so as to besimilar to that in FIG. 6. Utilization of aircraft direct lift controlto achieve proper operational configuration as illustrated in FIGS. 1-7is a major feature of the present invention. In one embodiment this isachieved by apparatus of the type illustrated in FIG. '8, but using theangle sensors of FIGS. 9 and 12, as described hereinafter. In FIG. 8,the overall physical dynamics of each aircraft as a body, including itsaltitude, position, acceleration and mass, are represented by the blocks22', respectively. Apparatus associated with each aircraft is located onthe same side of FIG. 8 therewith. In accordance with one aspect of thepresent invention, the lead aircraft is operative under manual control,the pilot thereof executing such maneuvers as are called for in a flightplan in the task being achieved in the dual lift mode. The pilotoperates the collective'pitch control 46 to induce direct lift (up, inthe frame of reference to the aircraft). This control is connected bymechanical means 48 (such as push rods, bell cranks or linkage, etc.) tooperate the collective pitch adjustment, to change the pitch of all therotor blades simultaneously so as to cause the aircraft to experiencedirect lift. The collective pitch control 46 is also directly connectedto a position sensor 50 which derives a proportional electrical signalwhich is conducted by a line 52 downwardly along cable 24, across thebar 32 and upwardly along the cable 26 (both schematically shown otherwise in FIG. 8) and emphasized herein by looping 52' in the line 52, toa summing network 54 in the trailing aircraft. With the aircraft flyingin perfect formation, the signal on line 52 is passed through thesumming network 54, with no other input, to an outer loop eIectro'mechanical actuator 56 which is connected by mechanical means 58 tooperate the collective pitch control 60 of the trailing aircraft. It isthis control which is mechanically linked by similar means 62 to all ofthe rotor pitch mechanisms of the trailing aircraft so as to inducedirect lift in the trailing aircraft (22'). This form of control ofcollective pitch on the trailing aircraft allows the trailing aircraftpilot to override the inputs from the leading aircraft, if he shoulddeem it necessary under any given set of circumstances.

The control just described is adequate to accommodate the situation of adirect lift induced in the leading aircraft to induce a correspondingdirect lift in the trailing aircraft. However, an additional input tothe summing network 54 is required for correction of faulty flightconfigurations (as in FIGS. 5 and 7 hereinbefore).

To achieve this, the bar 32 has mounted thereon a cable angle sensor 64,which may comprise any suitable means of sensing a displacement of thecable 26 from the normal to the plane in which the bar 32 lies. Forinstance, the cable angle sensor 64 may comprise a simple potentiometerarrangement as is known in the art (and utilized widely for similarsituations, such as sonar buoy hovering operations). The cable anglesensor 64 provides an electrical signal manifestation of the angle on aline 66 for application to a function generator 68 which provides aninput to the summing network 54 over a line 70. The function of cableangle (Ks K K/s) includes the angle (K) its rate of change, Ks (where sis the LaPlacian operator equivalent of d/dt), and its integral (K/s).The constants (K) are amplification factors, and are described morefully hereinafter. The Ks term is for damping; and K term is positionalto drive the trailing aircraft 22' to a null situation where its cable26 perpendicular to the bar 32; the K/s term provides long term responseto even minute amounts of drift or other long term errors in the system,as well as providing an input in the case where there is a long range.direct lift change in the lead aircraft to assist in providingadditional lift to the trailing aircraft to help close the gap.

Thus, the direct lift control illustrated in FIG. 8 provides direct liftto the trailing helicopter, so as to maintain the cable 26 in a positionwhich is perpendicular with respect to the bar 32. Note that, in theembodiment of FIG. 8, all control is supplied solely to the trail ingaircraft, and only the angle between the trailing cable 26 and the bar32 is sensed, as disclosed and claimed in a copending application of thesame assignee, filed on even date herewith (Aug. 2, I971) by J. R.Maciolelt, Ser. No. 168,200. This embodiment is ideal in a multi-liftcontrol system employing all the controls (or comparable controls) whichare disclosed herein. On the other hand, this aspect of the presentinvention may be employed in a partial multi-lift control system havingonly direct lift control, and relying upon manual control for otherpositioning of the aircraft. In such a case, the embodiment partiallyillustrated in FIG. 9 may be preferable. Therein, an additional sensor72 is provided for the forward cable 24 to sense the angle of the cable24 with respect to the leading end of the bar 32, and the difl'erencebetween the two angles is taken in a summing circuit 74 so as to providea signal on a line 76 to a functional unit 78, which providesessentially the same function as the unit 68 of FIG. 8 (except perhapsfor a different constant to accommodate the fact that a different angleparameter is being utilized). By utilizing the difference in the angleas a direct lift parameter, a situation which may arise under manualcontrol as illustrated in FIG. would not result in a direct lift.Therein, the two aircraft 20, 22 are flying too close togetherlaterally, which results in an angle in each of the cables. Althoughthis is likely to occur to some degree under manual control, the extentthereof is exaggerated slightly in FIG. 10. However, in the situation ofFIG. 10, in straight line flight, direct lift of the trailing aircraft22 would result in the situation illustrated roughly in FIG. 11.However, this is not corrective, and in fact results in an excess of theload being placed upon the aircraft 22. On the other hand, since theangles are equal in the situation in FIG. 10, the aircraft merely flyingtoo close together, the embodiment in FIG. 9 would provide no directlift input to correct the situation, thereby giving the operator of thetrailing aircraft time to alter his general position with respect to thelead aircraft so as to correct the situation of FIG. 10 under manualcontrol. Still another embodiment of this aspect of the invention isillustrated in FIG. 12, wherein optical (orother radiation) detection ofthe angle of the lead aircraft with respect to the reference platform ofthe trailing aircraft, corrected for pitch and roll by the verticalgyro, provides substantially the same information as the embodiment ofFIG. 9. In FIG. 12, the lead aircraft 20 may be provided with a pulsedelectromagnetic radiation emitter 80 which operates a position sensitivedevice 82 on the trailing aircraft 22, the sensor 82 being positioned soas to supply a null when the angle between the two aircraft is thedesired anglgjpr the proper relative elevation and perpendicularrelationship with respect to the bar 32. The sensor 82 may provide asignal on the line 84 to a summing unit 86 which subtracts the effectsof pitch and roll supplied thereto as electrical signals on lines 88, 90from the vertical gyro 92 of the trailing aircraft. The output of thesumming unit 86 appearing on the line 76' may be utilized in the samefashion as illustrated in FIG. 9. The exact nature of the type ofsensing to be utilized may depend on such factors as choice between acomplete control system as is disclosed herein, and utilization ofdirect lift control only (with other positioning in the twin-liftoperation being achieved by manual control).

The particular constants which may be provided in the function generator78 of FIG. 9, are all determined in dependence upon a particularutilization of the present invention. The constants for all of thecircuits herein are determined by a stability analysis of the loadsystem in conjunction with the operating characteristic of the aircraftinvolved. Factors that vary the values of the constants which aredeterminable by a stability analysis of the load system are the lengthof the bar, the lengths of the cables, the relative positioning desiredfor the aircraft, and the load itself. Adjustability between constantsis desirable for various load situations. However, as described brieflyhereinbefore, it should be understood that a wide range of loadsituations would not normally be anticipated by a twin-lift system ofthe type disclosed herein due to the fact that loads less than half ofthe maximum load achievable by both aircraft would normally be handledby a single aircraft because of the inherent simplicity, safety andreliability of a single lift situation. Therefore, the range of duallift load is from maximum to roughly half of maximum, and the incrementstherein may be fairly broad so that a wide range of constants is notrequired.

In FIG. 13, the lateral or roll cyclic pitch controls for both the leadaircraft 20' and the trailing aircraft 22' include modifications to theinner or stability control loop. The trailing aircraft has modificationsto the outer or general control loop. The controls normally found ondirect lift aircraft (such as helicopters) are illustrated within dottedlines designated AFCS (automatic flight control system). Lateral cyclicpitch controls of this type, as well as the longitudinal and collectivepitch and other basic controls for a helicopter, are illustrated in anautomatic flight control system disclosed in U. S. Pat. No. 3,071,355,FLIGHT CON- TROL SYSTEM, issued to E. S. Carter, Jr. on Jan. 1, 1963 andassigned to a common assignee herewith. As described hereinbefore, thepilot controls the lead aircraft, subject only to inner loop stabilitycontrols and inner loop controls relating to the present invention. Thepilot, in operating a four-way cyclic pitch control stick, can controllateral and longitudinal motion of the aircraft by inducing cyclic pitchinto the rotor blades. That portion of the control stick input whichrelates to cyclic pitch control for lateral motion is designated as thelateral cyclic pitch control 100. Movement of the control is combinedwith a mechanical input from an electromagnetic actuator 102 in anysuitable mechanical motion combining means 104 (which may comprise anysuitable form of pushrods, bell cranks, linkages, and so forth), tocontrol the cyclic pitch of the rotors within the aircraft 20' so as tocause lateral tilting of the aircraft to port or starboard. All of thisis well known in the art as exemplified by the aforementioned Carterpatent. In the description of FIG. 13, the aircraft is assumed to be inan orthogonal coordinate system with the X coordinate equal to theheading of the aircraft, the Y coordinate being in the plane with the Xcoordinate but lateral (to starboard), and the Z coordinateperpendicular to the X-Y plane and downward with respect to theaircraft. Thus, roll of the aircraft induces lateral motion in the Ydirection. In a basic automatic flight control system of the type knownto the art, the electromagnetic actuator 102 responds directly to afunction generator 106 which provides the rate of roll as sensed by theroll gyro 108 of the aircraft, which is applied in an opposite sense, bythe electromagnetic actuator 102 to the mechanical motion of the lateralcyclic pitch control 100 so as to counteract the manual motion inducedby the pilot, thereby to cause the aircraft to roll at nearly constantrate. This is a form of stability normally provided by the inner loop ofaircraft of this type in accordance with the teachings of the prior art.

In accordance with the present invention, motion of the lateral cyclicpitch control 100 is also coupled to a position sensor 110 to generatean electrical signal on a line 1 12 which is a manifestation of theposition of the control 100. The line 112, similar to the line 52 inFIG. 8, extends down the cable 24, across the bar 32 and up the cable 26to the trail aircraft, as emphasized herein by the loop 112' in the line112. The signal on the line 112 is amplified by an amplifier 114 andapplied over a line 115 to an electrical summing unit 116, the output ofwhich is converted by an electromechanical actuator 118 to drive thelateral cyclic pitch control 120 of the trailing aircraft. The control120, except for the fact that it can be driven by the electromechanicalactuator 118, is the same as the control 100 and others known in theprior art. Its output is a mechanical motion which has combinedtherewith, in a suitable means 122, a me chanical input from anelectromechanical actuator 124 which responds to the inner loop orstability portion of the automatic flight control system of the trailingaircraft. In manual operation of an aircraft, the pilot controls theaircraft affirmatively while viewing instruments to cause the aircraftto return to a horizontal position after executing turns and orelevation changes. In order to permit hands-off flying of the trailingaircraft in accordance with the present invention, the outer or primarycontrol loop of the trailing aircraft receives, at the summing unit 116,an electrical signal on a line 126 indicative of the absolute rollattitude of the aircraft as developed by a roll gyro 128. This is thesame sort of roll gyro as the roll gyro 108 and as utilized in theautomatic flight control systems of aircraft of this type known to theprior art. Although not illustrated herein, the lead aircraft may alsohave gyro inputs to cyclic pitch to allow automatic control overhorizontal position. The roll gyro 128 will provide no signal whatsoeverwhen the trail aircraft is perferctly horizontal with respect to theearth; however, if for some reason the aircraft gets out of thehorizontal position (without a strong signal on the line 115 causing itto do so) the sense of the signal on the line 126 is such as will causethe electromagnetic actuator 118 to vary the lateral cyclic pitchcontrol 120 so as to return the aircraft to an upright position.Although rate control is normally provided in the loop of the automaticflight control system, the rate control would stop the roll, but notcorrect it; that is why the additional input from the roll gyro to theouter loop summing unit 116 is required. The rate of roll is applied byan amplifier 130 to the electromagnetic actuator 124 so as to supplyroll stability as described with respect to the lead aircraft,hereinbefore. The summing unit 116 also receives information from FIG.16, which is a function of the deviation of the trail aircraft in its Ydirection from its correct position in the desired formation(asillustrated in FIG. 3). The nature and function of this input isdescribed more fully with respect to FIG. 16, hereinafter; its effect onthe lateral cyclic controls, however, is to cause the change in cyclicpitch so as to move the aircraft in the Y direction to reposition it ina correct position in the desired formation.

Thus there has been described with respect to FIG. 3 thus far, thelateral cyclic pitch controls which comprise the basic inner loop forboth the leading and the trailing aircraft, and the outer control loopof the trailing aircraft being slaved to the manual control of the leadaircraft.

In each of the aircraft 20, 22, the electromagnetic actuators 102, 124in the inner loop of the lateral cyclic controls, receive inputs fromrespective summing units 132, 134 which include inputs in addition tothose representing roll gyro rate from the amplifiers 106, 130. Theother inputs to the summing units 132, 134 relate to positionalstability and stability of the cables 24, 26. Specifically, accelerationin the Y direction is sensed by Y accelerometers 136, 138, the signalsfrom which are applied to respective summing units 140, 142. Because ofthe fact, as is known in the art, that any roll of the aircraft willprovide a small gravitational component in the Y direction of theinertial reference platform of the aircraft, the summing units 140, 142each receives inputs relating to the sine of the roll angle fromrespective functional units 144, 146 which subtract, from the Yacceleration, a value equal to gravity times the sine of the roll angle.The resulting output from each of the summing units 140, 142 comprises asignal manifestation of true Y component of acceleration in the inertialreference plane of the aircraft. These signals are applied to respectivelow pass filters 148, 150 which filter out high frequency noise,vibrations, and so forth. The output of the filters 148, 150 are appliedto respective summing units 152, 154 so as to apply to the summing units132, 134, negative signals which mitigate against lateral cyclic pitchin an amount proportional to the acceleration of the aircraft in the Ydirection. Thus, the outputs of the filters 148, 150, provide to eachaircraft, through its respective inner loop, stability against lateralacceleration.

Stability of the load itself is also provided for each aircraft in FIG.13. In each aircraft, the related cable 24, 26 preferably extendsdownwardly out of the aircraft from a cable winch (of the type known tothe prior art). In doing so, the cable passes a pair of spring loadedcable position sensors, one oriented in the X-Z plane so as to sensemotion of the cable in the X direction, and one oriented in the Y-Zplane so as to sense motion in the Y direction. The sensing arms may beconnected to potentiometers to generate electric signals proportional toangles of the cable in the respective plane; or, other suitable cablesensing means may be utilized. In FIG. 13, each aircraft includes a Ydirec tion cable angle sensor 156, 158 for sensing the angle of therelated cable 24, 26 with respect to the Y,Z plane only. Signalsindicative of the respective angle are applied over related lines 160,162 to a related series of functional blocks 164-169. The blocks 164,165 provide signals to a pair of summing units 170, 172 which are afunction of the rate of change of the angle, such signals being utilizedfor damping purposes in a well known fashion. This serves to preventpendular motion of the load. The functional blocks 166, 167 providesignals to related summing units 174, 176 which are a function of theabsolute angle of the related cable with respect to the aircraft, whichsignals are utilized to cause the aircraft to move laterally so as tostraighten up the cable in the lateral plane, thereby tending tomaintain the cable perpendicular to the related aircraft. The functionalblocks 168, 169 employ the same constant as the related one of thefunctional blocks 166, 167 but provide the angle-limited integral of theangle; these are log filters, with a time constant of 11". Thus, thefunctional blocks 168, 169 initially provide no output, but afterseveral time constants, provide exactly the same output as the relatedfunctional blocks 166, 167, as disclosed and claimed in theaforementioned copending application. Thus, if there is an additionalchange in the position of the related cable 24, 26 with respect to therelated aircraft reference frame 20', 22', there is a commensurateoutput from the related amplifier 166, 167 which tends to cause therelated aircraft to move laterally to remove the lateral angle from thecable 24, 26. However, if the angle is not removed, in a matter ofseconds, then this fact is indicative of a maneuvering of the aircraftwhich requires that there be a net lateral angle between the aircraftitself and the cable 24, 26 related thereto. Such a case occurs in thecase of making a turn, as is illustrated in FIG. 6 hereinbefore.Although it appears in FIG. 6 that the entire formation has been movedinto an angle with respect to the horizontal, in fact, the aircraft aremore nearly horizontal than is the remainder of the formation (such asthe spreader bar 32). Thus, to maintain the formation of FIG. 6, theaircraft will be at a slight lateral angle with respect to theirrespective cables, since as the turn is made, the aircraft commence theturn and the load always lags behind the aircraft in making the turn dueto drag imposed thereon and inertia. This small angle is maintained bythe fact that the limited integral units 168, 169 cancel out theamplifier units 166, 167 after several time constants, to allow thecables to seek their own level in a maneuver. The small actual lag isequivalent to the lag built into the limited integral functional units168, 169.

The manner of arranging the functional blocks and summing units in FIG.13 may be altered to provide the same results in simpler configuration,if desired.

The longitudinal or pitch cyclic pitch controls illustrated in FIG. 14are identical to the lateral or roll cyclic pitch controls illustratedin FIG. 13, with the exception that the values of the constants may bedifferent and that all of theparameters therein relate to thelongitudinal, or pitch motion, and to the X direction in thereference'frame of the aircraft. Additionally, the limited integrators178, 180 at the bottom of FIG. 14 provide for drag during accelerationsand in forward flight, as illustrated in FIG. 4.

Although it is relatively easy for a pilot to maintain heading of thetrailing aircraft substantially the same as the heading of the leadaircraft, in order to maintain a stable positioning of the load withrespect to the aircraft while in a stable pattern or formation, thepresent invention also provides automatic control over heading (or yaw)of the trailing aircraft as illustrated in FIG. 15. Therein, a rotaryrudder pedal 182, which varies the collective pitch of the rotary rudderto alter the heading control for the lead aircraft 20, is mechanicallyconnected to a position sensor 184 which provides a signal on a line 186indicative of pilot-induced motion of the rotary rudder pedal 182. Theline 186 extends downwardly on the cable 24 across the bar 32 andupwardly on the cable 26, as is illustrated herein by the loop 186' inthe line 186, and is passed to an amplifier 190 in the trailing aircraftfor application to a summing unit 192.

The summing unit 192 operates an electromechanical actuator 194 so as tooperate the rotary rudder pedal 196 of the trailing aircraft, the outputof which is mechanically coupled to the rotary rudder (or tail rotor) ofthe trailing aircraft Thus, motion of the rudder control 182 in the leadaircraft induces a similar motion in the rudder control 196 of thetrailing aircraft. Each of the aircraft 20, 22 are provided with rotaryrudder inner loop controls, which utilize an electric signal from arelated gyro compass 198, 200, indicative of the heading of theaircraft, passed through a differentiator 202, 204 to anelectromechanical actuator 206, 208 to provide a mechanical input tomotion combining means 210, 212 so as to stabilize changes in headinginduced by motion of the rudder pedal 182, 196.

Herein, the letter W designates angle, such as heading, with suitablesubscripts to indicate heading of the lead aircraft, the trailingaircraft, the bar, the error between the lead aircraft heading and thebar heading, the desired primary angle of the formation, etc. Thus, theheading of the lead aircraft is designated in FIG. 3 as W L the trailingaircraft as W and the bar as W and the error between the lead aircraftand the trailing aircraft as W In the heading controls of FIG. 15, theheading of the trailing aircraft (W on a line 213 is subtracted from theheading of the lead aircraft (W,,) on a line 214 by a summing unit 215.The line 214 extends down the cable 24, across the bar 32, and up thecable 26 as indicated by the loop 214' in the line 214. The differencein headings (W is applied to a functional block 216 (which is identicalto the functional block 68 in FIG. 8, except that different constantsmay be utilized therein). This block generates an output signal whichcomprises absolute difference in heading, the rate of change ofdifference of heading, and the accumulation of heading differences byutilizing the Ks, K, K/s functions as described hereinbefore. The outputof the block 216 is applied over a line 217 to the summing unit 192 soas to provide corrective outer loop control over the heading of thetrailing aircraft in response to the actual heading of the leadaircraft. Thus, the outer loop heading control of the trailing aircraftis responsive not only to commands made by the pilot of the leadingaircraft to change heading, but in accordance with the invention of saidcopending application, differences in the response of the two aircraftare accommodated by functions of the difference in actual headingbetween the two aircraft.

It should be noted that in the case of the heading controls of FIG. 15,there is no modification to the inner loop of either aircraft; thechanges required for this invention are simply command (and relativeheading, if desired) control over the outer loop of the trailingaircraft in responseto commands in (and actual relative heading of, ifdesired) the lead aircraft;

Maintenance of a desired formation is assisted, in accordance with thisinvention, by a means of position controls illustrated in FIG. 16. Thesecontrols are responsive to heading signals which are derived in FIG. 15.A summing unit 220 responds to the lead aircraft heading (W from thegryo compass 198, and the bar heading (W derived from a gyro compass 222mounted directly on the spreader bar 32. The gyro compass 222 is somounted with respect to the bar 32 that a relative headingof zero isdefined as being perpendicular to the longitudinal axis of the bar. Bydoing this, as is described more fully hereinafter, it is possible toalter the primary angle (W FIG. 3) of a desired formation, quite simply,without altering the position of the gryo 222 on the bar 32. The bargyro compass 222 is connected by wiring 223 which extends along the barto the cable 24 and upwardly along the cable 24 to the lead aircraft 22.

Referring now to FIG. 3, a desired formation is therein illustrated ashaving the bar 32 oriented at about 30 with respect to the heading ofthe formation. The heading (W,,) of the lead aircraft 20, the motionheading (W.,) of the bar 32 and the heading (W of the trailing aircraft22 are all equal, when in the desired formation, without turns. Theactual heading of the bar (W differs from the direction of motion of thebar (W,,) by the desired primary angle of the formation (W,,). In otherwords, when in proper formation, the motion of the bar (W equals theheading of the bar (W plus the desired formation angle (W The desireddistances in the X and Y direction (X,,, Y,,) are related to the lengthof the bar (L) and the desired primary angle of the formation (W,,) asfollows:

X,, LcosW AND Y Lsin W To find the X and Y errors in the formation, sothat these errors can be utilized in the longitudinal and lateral cycliccontrols, respectively, to correct the formation to that which isdesired, consider first the error, X in the X direction:

X X-X Lcos( W W LcosW Lcos( W W W LcosW X; Lco s[( W W W ]LeosWSimilarly, the error in the Y direction is given as:

Y,,- Lsin[( W W W ]LsinW,,

In FIG. 16, a signal indicative of the error in the X direction (X isprovided on a line 230, and a signal indicative of the error in the Ydirection (Y is provided on the line 232. For each of the equations (7)and (8), the term (W W is provided on a line 234 to a summing unit 236which subtracts the term W which is provided thereto on the line 238from a position sensor 240 which is mechanically adjusted by a W,,control input 242. In other words, primary angle desired for theformation is manually adjusted by adjusting a knob on a potentiometer orother suitable position sensor 240 so as to supply a desired angle inputto the summing network 236. The net term [(W W W is thus provided by thesumming unit 236 on the line 244. Then, a pair of functional units 246,248 provide L times the cosine and sine, respectively, of this net angleto related summing units 250, 252. Similarly, identical functional units254, 256 provide L times the cosine and sine of the desired angle (W onthe line 238 to the summing units 250, 252, thereby to derive the X andY errors in accordance with equations (7) and (8) as describedhereinbefore. The X and Y errors on the lines 230, 232 are converted todamped positional errors by related functional blocks 258, 260: in eachof the units 258, 260 the Ks term will immediately create an input tostop the change of error in X or Y and the K tenn will tend to removethe error once the change has been eliminated. These signals areapplied, in accordance with this invention and as describedhereinbefore, to FIGS. 14 and 13, respectively, so as to cause direct,outer loop inputs to the trailing helicopter to tend to maintain correctX and Y positioning of the trailing aircraft in the desired formationillustrated in FIG. 3.

Although illustrated herein with the summing unit 220 and the positioncontrols of FIG. 6 being disposed in the lead aircraft 20, they couldobviously be equally well disposed in the trailing aircraft 22.

The embodiments of the invention described hereinbefore can besummarized, as illustrated in FIGS. 17 and 18. FIG. 17 illustrates thewell known control loops of a single aircraft of a type which may beutilized in the system in acordance with this invention. In FIG. 17, theaircraft 280 is controlled by manual controls 282 in response tomovement thereof by thepilot 284. However, the controls are modified bymechanical inputs provided thereto by an electromagnetic actuator 286which responds to the inertial system 288 of the aircraft. Thus, aninner control loop includes modifications to pilot movements. The outeror general control loop includes the pilot, the manual controls 282, theactual changes in the control surfaces of the aircraft 280, the responseof the aircraft (designated by the block 290) as well as the pilotsobservation of the response; the action of the pilot, in response to hisobservations, closes the outer loop of the control system.

One aspect of the present invention is easily compared directly with theinner and outer loops of the known aircraft in FIG. 17. In FIG. 18, thelead aircraft 20' has the same outer loop including the aircraftresponse 290, the pilot 284 and the manual controls 282.

However, the outer loop of the trailing aircraft differs significantlysince it comprises simply a response to the motion of the lead pilot 284as a result of position sensors 292 electrically connected by wiring 294down the cable 24, across the load 32 and up the cable 26 (as indicatedby 294) to electromagnetic actuators 296 which operate respective manualcontrols 298 in the trailing aircraft 22'. The trailing aircraft pilot300 may override the manual controls 298 if desired. The outer loop ofthe trail aircraft also includes functions of relative heading of thelead aircraft with the load and inertial inputs to maintain the trailingaircraft in a correct horizontal position in a desired formation. Anadditional aspect of the present invention is that each of the aircraftinclude, in addition to the original inner loop I controls (such asillustrated in FIG. 17), further inner loop controls in response to theinertial system 302, 304 of the respective aircraft and in response toangle sensors 306, 308 relating to angles between the cables extendingbetween the aircraft and the spreader bar 32. The additions to therespective inner loops provide, with respect to each aircraft, formationstability and load stability.

Thus, in accordance with the present invention, an outer control loopfor an aircraft includes control functions over the physical positioningof the manually movable controls utilized by the pilot of an aircraft toinduce responses therein; an inner loop control system includes controlinputs utilized in motion combining means to alter the actual motion ofcontrol surfaces of the aircraft from those that would be caused by themotion of the manual controls, in amounts which relate to stabilizingthe aircraft, both in its inertial reference and in its relationship toa load being supported by a plurality of said aircraft. One aircraft isthe lead or master aircraft and its outer control loop is unmodified,relying solely on the pilot for the operation of the manual ,controlsthereof, but its inner control loop modifying the motions of controlsurfaces induced by the manually movable controls as a result ofinertial, positional, and relative load position stabilizing inputs. Thetrailing or slave aircraft, on the other hand, has motion of the firstaircraft coupled into its outer control loop, so that the pilot thereof,in the absense of an override due to special circumstances, observes themanually movable controls moving automatically in response to motionsimparted thereto by actuators. In addition to coupling motion of thelead aircraft controls into the trailing aircraft controls, the trailingaircraft outer loop controls include pitch and roll stabilization, aswell as direct lift to maintain a proper altitude with respect to theload. The trailing aircraft has the same inner loop inputs forstabilizing the absolute position of the aircraft and its relationshipto the load, as does the leading aircraft. Thus, the outer loop of thetrailing or slave aircraft includes only the slower, general positionalinputs of the type usually imparted thereto by the pilot, and the innerloop includes the more rapid stabilizing inputs which the pilot normallyis not able to compensate for and does not wish to sense in the stick.

The embodiments herein have been described primarily with respect tohelicopters of the single main rotor type, having a main rotor capableof cyclic and collective pitch and having a rotary rudder. However,

it should be obvious from the functional descriptions herein, that thecontrols herein, modified to suit the given characteristics of anaircraft, may be utilized for multiple-aircraft lifting of loads bymeans of aircraft other than helicopters. Additionally, the embodimentherein is described with respect to a dual lift helicopter operation,but it should be understood that the system herein may be employed inoperations involving more than one trailing or slave aircraft.The'manner in which the utilization of more than one slave aircraftmodifies the responses, required in other aircraft is accommodated inthe stability analysis of thetotal system including as many aircraft asdesired. This naturally becomes more complex for anygreater number ofaircraft; however, the principles are the same and the analyses to beutilized are the same as for the dual-lift case, which is disclosedherein only for simplicity. Also, the aircraft may be coupled bywireless means, such as used in telemetry.

Thus, although a multiple lift aircraft control system in accordancewith the present invention has been described with respect to preferredembodiments thereof, it should be understood by those skilled in the artthat the foregoing and various other changes and omissions in the formand detail thereof may be made therein without departing from the spiritand the scope of the invention.

Having thus described typical embodiments of our invention, that whichwe claim as new and desire to semeans coupling themanual controls of afirst one of said aircraft to automatically move the manual controls ofa second one of said aircraft in response to motion imparted to saidfirst aircraft manual controls;

means disposed in each aircraft for sensing a positional relationship ofsaid common load to the related one of said aircraft; and

signal means in each of said aircraft responsive to the related one ofsaid positional relationship sensing means and connected to the relatedone of said inner loop control systems for generating signals operativethrough said inner loop control system to modify the related outer loopcontrol system in a manner to tend to stabilize the positionalrelationship of the related one of said aircraft with respect to saidcommon load. 2. The system according to claim 1 wherein: said load issupported by support cables extending generally downward from each ofsaid aircraft;

said positional sensing means comprise means for sensing the relativeangle in the lateral direction and the relative angle inthe longitudinaldirection of the related one of said support cables with respect to thereference platform of the related aircraft;

and said signal means comprise means responsive to respective ones ofsaid sensing means to respectively modify the lateral and longitudinalouter loop control system of said aircraft in response thereto. 1

3. The system according to claim 2 wherein said modification is inresponse to at least the rate 'of change of the related angle.

4. A system according to'claim 1 additionally comprising meas in each ofsaid aircraft responsive to the related one of said inertial systems andconnected to the related one of said inner loop control systems forstabilizing the related outer loop control system in a manner tending tostabilize the position of the related aircraft.

S. The system according to claim 1 wherein said outer loop controlsystem of each aircraft includes con-' trol over longitudinal cyclicpitch, lateral cyclic pitch and collective pitch of a main rotor, andincludes co]- lective pitch of a rotary rudder.

6. The system according to claim 5 wherein said inner loop controlsystem of each aircraft includes stabilization in respect of pitch rate,rollrate, lateral acceleration, longitudinal acceleration, and rate ofheading change.

7. The system according to claim I wherein the outer control loop ofsaid second aircraft includes pitch and roll inputs from the inertialsystem of said second aircraft, thereby providing control over thehorizontal stability of said one aircraft.

8. The system according to claim l additionally comprising:

an inertial system on said load;

and means connected to the inertial systems of said first aircraft andsaid load and responsive to the difference between the heading of saidfirst aircraft and the heading of said load for modifying the outer loopcontrol system of said second aircraft to tend to maintain the relativeposition of said second aircraft with respect to said first aircraft aswill provide a desired relative heading of said load with respect tosaid aircraft, in a desired formation.

9. A multiple lift aircraft control system, comprising:

a plurality of aircraft;

a support cable extending generally downwardly from each of saidaircraft;

a spreader bar connected, at a related one of a plurality of points, tothe downmost end of each of said cables;

a load suspended from said spreader bar in a manner to load therespective points of said spreader bar substantially equally;

a plurality of means, one mounted on a first one of said aircraft andone mounted on said spreader bar, each providing a signal manifestationof the heading of the body on which it is mounted;

and means responsive to the signal manifestations provided by saidheading means for controlling the lateral and longitudinal motion of asecond one of said aircraft, thereby tending to control the position ofsaid second aircraft to maintain said second aircraft in a correctposition in a desired formation.

10. A multiple aircraft load lift system comprising:

a plurality of aircraft, at least a first one of said aircraft having adirect lift capability and a related control system therefor;

a support cable extending downward from each of said aircraft;

a spreader bar connected to the downward end of each of said supportcables;

means for sensing a first angle between said spreader barand the one ofsaid support cables relating to said first aircraft;

means for sensing a second angle between a second one of said supportcables and said spreader bar;

means responsive to said angle sensing means to generate a signal inresponse to the difference between said first and second angles; and

lift means responsive to said difference signal means for inducing achange in the direct lift of one of said aircraft in response to saiddifference signal, said change in direct lift causing motion of said oneaircraft in a manner to tend to drive said first angle to 11. A multipleaircraft load lift system comprising:

a plurality of aircraft, at leasta first one of said aircraft having adirect lift capability and a related control system therefor, and anattitude reference platform;

a support cable extending downwardly from each of said aircraft;

a spreader bar connected to the downward end of each of said supportcables;

means for sensing the angle in the vertical between said referenceplatform and a second one of said aircraft and for generating a signalindicative thereof;

lift means responsive to said angle signal means for inducing a changein the direct lift of said first aircraft in response to said anglesignal, said change in direct lift causing motion of said first aircraftin a manner to tend to drive said angle to a desired angle 12. A controlsystem for controlling a plurality of aircraft supporting a common loadsupported by a spreader bar, said spreader bar having a point relatingto each of said aircraft, each of said aircraft including a supportcable extending generally downwardly therefrom and attached to therelated point on said spreader bar, each aircraft including an outerloop control system in which motion of manually movable controls alterthe position of control surfaces of the aircraft causing positionalresponse in the aircraft including direct lift response, comprising:

a plurality of electromagnetic position sensors disposed in a first oneof said aircraft, each sensor responsive to a motion of said manuallymovable controls relating to a given positional response of saidaircraft, each for generating an electrical signal manifestation of therelated motion; plurality of electrical conductors extending from saidfirst aircraft to a second one of said aircraft, each relating to acorresponding one of said signal manifestations;

' a plurality of electromechanical actuators disposed in said secondaircraft, each connected to a respective one of said electricalconductors, each disposed relative to the manual removable controls ofsaid second aircraft and responsive to a given signal manifestation inthe related conductor so as to cause motion therein to induce a responsein said second aircraft of the same type and substantially the samemagnitude as that in said first aircraft which causes said given signal,whereby manual operation of said first aircraft is electromechanicallycoupled to correspondingly operate said second aircraft;

a cable angle sensing means disposed on each of said aircraft to sensethe angle between said spreader bar and the related one of said cables,each of said angle sensing means developing an electrical signalmanifestation of said angle;

difference means responsive to said angle sensing means for generating adifference signal in response to the difference in angle represented bysaid angle signal manifestations; and

electrical signal combining means connected to said difference means andconnected in series with the one of said electrical conducting meansrelating to said direct lift response and jointly responsive to signalsthereon and to said difference signal to operate the related one of saidelectromechanical actuators, whereby direct lift in said second aircraftis controlled not only in response to direct control motion in saidfirst aircraft, but also in response to the relative angles between thesupport cables of said aircraft and said spreader bar.

1. A control system for controlling a plurality of aircraft supporting acommon load, each aircraft including an inertial system and an outerloop control system in which motion of manually movable controls altersthe position of control surfaces of the aircraft causing positionalresponse in the aircraft, and each including an inner loop controlsystem responsive to the related inertial system for automaticallystabilizing the related outer loop control system, comprising: mEanscoupling the manual controls of a first one of said aircraft toautomatically move the manual controls of a second one of said aircraftin response to motion imparted to said first aircraft manual controls;means disposed in each aircraft for sensing a positional relationship ofsaid common load to the related one of said aircraft; and signal meansin each of said aircraft responsive to the related one of saidpositional relationship sensing means and connected to the related oneof said inner loop control systems for generating signals operativethrough said inner loop control system to modify the related outer loopcontrol system in a manner to tend to stabilize the positionalrelationship of the related one of said aircraft with respect to saidcommon load.
 2. The system according to claim 1 wherein: said load issupported by support cables extending generally downward from each ofsaid aircraft; said positional sensing means comprise means for sensingthe relative angle in the lateral direction and the relative angle inthe longitudinal direction of the related one of said support cableswith respect to the reference platform of the related aircraft; and saidsignal means comprise means responsive to respective ones of saidsensing means to respectively modify the lateral and longitudinal outerloop control system of said aircraft in response thereto.
 3. The systemaccording to claim 2 wherein said modification is in response to atleast the rate of change of the related angle.
 4. A system according toclaim 1 additionally comprising meas in each of said aircraft responsiveto the related one of said inertial systems and connected to the relatedone of said inner loop control systems for stabilizing the related outerloop control system in a manner tending to stabilize the position of therelated aircraft.
 5. The system according to claim 1 wherein said outerloop control system of each aircraft includes control over longitudinalcyclic pitch, lateral cyclic pitch and collective pitch of a main rotor,and includes collective pitch of a rotary rudder.
 6. The systemaccording to claim 5 wherein said inner loop control system of eachaircraft includes stabilization in respect of pitch rate, roll rate,lateral acceleration, longitudinal acceleration, and rate of headingchange.
 7. The system according to claim 1 wherein the outer controlloop of said second aircraft includes pitch and roll inputs from theinertial system of said second aircraft, thereby providing control overthe horizontal stability of said one aircraft.
 8. The system accordingto claim 1 additionally comprising: an inertial system on said load; andmeans connected to the inertial systems of said first aircraft and saidload and responsive to the difference between the heading of said firstaircraft and the heading of said load for modifying the outer loopcontrol system of said second aircraft to tend to maintain the relativeposition of said second aircraft with respect to said first aircraft aswill provide a desired relative heading of said load with respect tosaid aircraft, in a desired formation.
 9. A multiple lift aircraftcontrol system, comprising: a plurality of aircraft; a support cableextending generally downwardly from each of said aircraft; a spreaderbar connected, at a related one of a plurality of points, to thedownmost end of each of said cables; a load suspended from said spreaderbar in a manner to load the respective points of said spreader barsubstantially equally; a plurality of means, one mounted on a first oneof said aircraft and one mounted on said spreader bar, each providing asignal manifestation of the heading of the body on which it is mounted;and means responsive to the signal manifestations provided by saidheading means for controlling the lateral and longitudinal motion of asecond one of said aircraft, thereby tending to control the position ofsaid second aircraft to maintAin said second aircraft in a correctposition in a desired formation.
 10. A multiple aircraft load liftsystem comprising: a plurality of aircraft, at least a first one of saidaircraft having a direct lift capability and a related control systemtherefor; a support cable extending downward from each of said aircraft;a spreader bar connected to the downward end of each of said supportcables; means for sensing a first angle between said spreader bar andthe one of said support cables relating to said first aircraft; meansfor sensing a second angle between a second one of said support cablesand said spreader bar; means responsive to said angle sensing means togenerate a signal in response to the difference between said first andsecond angles; and lift means responsive to said difference signal meansfor inducing a change in the direct lift of one of said aircraft inresponse to said difference signal, said change in direct lift causingmotion of said one aircraft in a manner to tend to drive said firstangle to 90*.
 11. A multiple aircraft load lift system comprising: aplurality of aircraft, at least a first one of said aircraft having adirect lift capability and a related control system therefor, and anattitude reference platform; a support cable extending downwardly fromeach of said aircraft; a spreader bar connected to the downward end ofeach of said support cables; means for sensing the angle in the verticalbetween said reference platform and a second one of said aircraft andfor generating a signal indicative thereof; lift means responsive tosaid angle signal means for inducing a change in the direct lift of saidfirst aircraft in response to said angle signal, said change in directlift causing motion of said first aircraft in a manner to tend to drivesaid angle to a desired angle.
 12. A control system for controlling aplurality of aircraft supporting a common load supported by a spreaderbar, said spreader bar having a point relating to each of said aircraft,each of said aircraft including a support cable extending generallydownwardly therefrom and attached to the related point on said spreaderbar, each aircraft including an outer loop control system in whichmotion of manually movable controls alter the position of controlsurfaces of the aircraft causing positional response in the aircraftincluding direct lift response, comprising: a plurality ofelectromagnetic position sensors disposed in a first one of saidaircraft, each sensor responsive to a motion of said manually movablecontrols relating to a given positional response of said aircraft, eachfor generating an electrical signal manifestation of the related motion;a plurality of electrical conductors extending from said first aircraftto a second one of said aircraft, each relating to a corresponding oneof said signal manifestations; a plurality of electromechanicalactuators disposed in said second aircraft, each connected to arespective one of said electrical conductors, each disposed relative tothe manual removable controls of said second aircraft and responsive toa given signal manifestation in the related conductor so as to causemotion therein to induce a response in said second aircraft of the sametype and substantially the same magnitude as that in said first aircraftwhich causes said given signal, whereby manual operation of said firstaircraft is electromechanically coupled to correspondingly operate saidsecond aircraft; a cable angle sensing means disposed on each of saidaircraft to sense the angle between said spreader bar and the relatedone of said cables, each of said angle sensing means developing anelectrical signal manifestation of said angle; difference meansresponsive to said angle sensing means for generating a differencesignal in response to the difference in angle represented by said anglesignal manifestations; and electrical signal combining meanS connectedto said difference means and connected in series with the one of saidelectrical conducting means relating to said direct lift response andjointly responsive to signals thereon and to said difference signal tooperate the related one of said electromechanical actuators, wherebydirect lift in said second aircraft is controlled not only in responseto direct control motion in said first aircraft, but also in response tothe relative angles between the support cables of said aircraft and saidspreader bar.